This invention relates to coatings for components exposed to high temperatures, such as the hostile thermal environment of a gas turbine engine. More particularly, this invention is directed to a protective coating for a thermal barrier coating on a gas turbine engine component, in which the protective coating is resistant to attack by contaminants present in the operating environment of a gas turbine engine.
Hot section components of gas turbine engines are often protected by a thermal barrier coating (TBC), which reduces the temperature of the underlying component substrate and thereby prolongs the service life of the component. Ceramic materials and particularly yttria-stabilized zirconia (YSZ) are widely used as TBC materials because of their high temperature capability, low thermal conductivity, and relative ease of deposition by plasma spraying, flame spraying and physical vapor deposition (PVD) techniques. Air plasma spraying (APS) is often preferred over other deposition processes due to relatively low equipment costs and ease of application and masking.
To be effective, TBC systems must strongly adhere to the component and remain adherent throughout many heating and cooling cycles. The latter requirement is particularly demanding due to the different coefficients of thermal expansion (CTE) between ceramic materials and the substrates they protect, which are typically superalloys though ceramic matrix composite (CMC) materials are also used. To promote adhesion and extend the service life of a TBC system, an oxidation-resistant bond coat is often employed. Bond coats are typically in the form of an overlay coating such as MCrAlX (where M is iron, cobalt and/or nickel, and X is yttrium or another rare earth element), or a diffusion aluminide coating. During the deposition of the ceramic TBC and subsequent exposures to high temperatures, such as during engine operation, these bond coats form a tightly adherent alumina (Al2O3) layer or scale that adheres the TBC to the bond coat.
The service life of a TBC system is typically limited by a spallation event brought on by thermal fatigue. In addition to the CTE mismatch between a ceramic TBC and a metallic substrate, spallation can be promoted as a result of the TBC being contaminated with compounds found within a gas turbine engine during its operation. A notable example is a mixture of several different compounds, typically calcia, magnesia, alumina and silica, referred to herein as CMAS. CMAS is a relatively low melting eutectic that when molten is able to infiltrate to the cooler subsurface regions of a TBC, where it resolidifies. During thermal cycling, the CTE mismatch between CMAS and the TBC (e.g., YSZ) promotes spallation, particularly TBC deposited by APS (hereinafter, xe2x80x9cAPSTBCxe2x80x9d) due to the ability of the molten CMAS to penetrate the porous structure typical of APSTBC. Various studies have been performed to find a TBC material that if deposited by APS is resistant to infiltration by CMAS, though none have been found sufficiently acceptable for production processes.
Accordingly, it would be desirable if the resistance of TBC to spallation attributable to CMAS infiltration could be improved.
The present invention generally provides a protective coating and method for protecting a thermal barrier coating (TBC) on a component used in a high-temperature environment, such as the hot section of a gas turbine engine. The invention is particularly directed to a protective coating that significantly reduces if not prevents the infiltration of CMAS into an underlying TBC.
The protective coating of this invention comprises alumina particles in a silica-containing matrix, and may be substantially homogeneous or formed of multiple layers having different compositions. The composition and relative amounts of alumina and matrix material in the protective coating enable the coating to react with molten CMAS, forming a compound with a melting temperature that is significantly higher than CMAS. As such, infiltration of molten CMAS into the TBC is significantly reduced or entirely avoided. Instead, the new compound solidifies on the surface of the protective coating, and either remains adherent or spalls as small fragments that can be safely exhausted through the turbine engine.
In addition to addressing the vulnerability of TBC to spallation from CMAS contamination, the protective coating of this invention has additional properties advantageous to gas turbine engine components. For example, the protective coating is more erosion resistant than conventional TBC materials such as YSZ, and can be readily deposited to have a smoother surface finish that improves the aerodynamic performance of the component. The composition of the protective coating can also be tailored to have reduced transmissivity to infrared radiation, thereby significantly reducing the heating of the component by thermal radiation. In view of these benefits, the present invention is able to significantly extend the life of gas turbine engine components protected by TBC.
Other objects and advantages of this invention will be better appreciated from the following detailed description.